System and method for fabricating a composite material assembly

ABSTRACT

A method for fabricating a composite material assembly includes: a) providing an assembly system, b) laying down a first module on a first mold, the first module comprising a first laminate covering a first laminate support structure, c) laying down a second module on a second mold, the second module comprising a second laminate covering a second laminate support structure and extending over the at least one removable insert, d) removing the at least one removable insert from the second mold, and e) assembling the first mold with the second mold while overlapping a section of the second laminate extending over the at least one removable insert over the first laminate.

FIELD OF THE INVENTION

The present invention generally relates to composite materials. Thepresent invention more specifically relates to a system and method forfabricating a composite material assembly.

BACKGROUND OF THE INVENTION

Composite material assembly, and more particularly fuselagemanufacturing through the use of multi-piece sections, typicallyrequires pre-solidification and cure of each piece prior to assemblethem with splices between individual sections or portions.

The limitations of this methodology are:

-   -   a minimum two-step cure is required;    -   additional mechanical fasteners are required at splicing joints        on primary structure components;    -   the methodology requires handling equipment and assembly jigs        (for out-of-mold operations);    -   long fuselage manufacturing time;    -   over thickness at joints resulting in stress concentration;    -   increases in weight of assembly; and    -   surface preparation is required prior to bonding.

Various solutions for assembly of multi-piece sections have beenproposed in the prior art.

U.S. Pat. No. 7,459,048 discloses a method of manufacturing a unitarysection of an aircraft fuselage including steps of disposing a thinlayup mandrel element onto the outer shell surface of a cylindricalinner mandrel shell to form a mandrel with a layup surface. The methodfurther includes steps of laying-up fibers onto the layup surface whilethe mandrel rotates to form a unitary pre-cured section of an aircraftfuselage.

WO 98/32589 discloses composite structures having a continuous skinformed using automated fiber placement methods. The multiple layers offibers are placed on a fiber placement tool including a mandrel bodysurrounded by a bladder. Uncured composite structures are created byplacing fibers around the fiber placement tool as discontinuous segmentsthat are capable of moving or sliding in relation to each other in orderto be expandable from within. The uncured structures are then expandedagainst the other surface of the molds by creating a vacuum between thebladder and the molds.

U.S. Pat. No. 7,325,771 discloses structures and methods for joiningcomposite fuselage sections using spliced joints attaching a firststiffener on a first composite part as well as a second stiffener on asecond composite part through a fitting. A strap is then used to splicethe first and second composite parts together.

US 2006/0251847 discloses a method of joining composite elements inwhich the bonding is done through the thickness of fiber compositelaminates in order to reduce interlaminar stresses usingnon-interlocking and interlocking bonds.

US 2009/0148647 discloses a method of fabricating composite structuresby joining a plurality of composite modules along their edges usingscarf joints instead of using advance fiber placement machines thatrequire high capital investment and operating costs.

However, there is still a need for a system and method for fabricatingcomposite material assemblies that facilitate assembly of parts whenforming structures while minimizing assembly equipment costs.

SUMMARY OF THE INVENTION

An object of the present invention is to propose a system and methodthat satisfies at least one of the above-mentioned needs.

According to the present invention, that object is accomplished with asystem for fabricating a composite material assembly comprising:

-   -   a first mold for receiving a first module made of composite        material, the first mold comprising:        -   a first composite material laminate support structure having            first and second opposite edges; and        -   a first attachment interface for attachment of the first            mold to an adjacent mold; and    -   a second mold for receiving a second module made of composite        material, the second mold comprising:        -   a second composite material laminate support structure            having first and second opposite edges;        -   a second attachment interface for attachment of the second            mold to the first mold; and        -   at least one removable insert extending beyond at least one            of the first and second edges of the second mold,            wherein the first module comprises a first laminate covering            the first laminate support structure, the second module            comprises a second laminate covering the second laminate            support structure and extending over the at least one            removable insert, and wherein the at least one removable            insert is removed from the second mold prior to assembly of            the first mold to the second mold, and a section of the            second laminate extending over the at least one removable            insert overlaps over the first laminate after closing and            assembly of the first mold onto the second mold.

According to the present invention, there is also provided a method forfabricating a composite material assembly comprising the steps of:

-   -   a) providing an assembly system comprising:        -   a first mold for receiving a first module made of composite            material, the first mold comprising:            -   a first composite material laminate support structure                having first and second opposite edges; and            -   a first attachment interface for attachment of the first                mold to an adjacent mold; and        -   a second mold for receiving a second module made of            composite material, the second mold comprising:            -   a second composite material laminate support structure                having first and second opposite edges;            -   a second attachment interface for attachment of the                second mold to the first mold; and            -   at least one removable insert extending beyond at least                one of the first and second edges of the second mold;    -   b) laying down the first module on the first mold, the first        module comprising a first laminate covering the first laminate        support structure;    -   c) laying down the second module on the second mold, the second        module comprising a second laminate covering the second laminate        support structure and extending over the at least one removable        insert;    -   d) removing the at least one removable insert from the second        mold; and    -   e) assembling the first mold to the second mold while        overlapping a section of the second laminate extending over the        at least one removable insert over the first laminate.

The present invention provides means for manufacturing one-piececomposite components originating from more than one mold while providinga structure that can be cured or solidified under heat and vacuum in onestep only, preferably with a composite material in a pre-prep form whichdoes not require autoclave treatment.

A non-restrictive description of a preferred embodiment of the inventionwill now be given with reference to the appended drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIGS. 1a to 1c are side cross-sectional views of the system according toa preferred embodiment of the present invention, showing an assemblysequence of a first monolithic laminate on a first mold onto a secondmonolithic laminate of a second mold with prior removal of a removableinsert;

FIGS. 2a to 2c are side cross-sectional views of the system according toanother preferred embodiment of the present invention, showing anassembly sequence of a first sandwich laminate on a first mold onto asecond sandwich laminate of a second mold with prior removal of aremovable insert, and a subsequent addition of a layup splice;

FIGS. 3a to 3e are front views of a build sequence of a tubularcomponent using the system according to another preferred embodiment ofthe present invention, using one removable insert per mold;

FIGS. 4a to 4e are perspective views of the build sequence of thetubular component shown in FIGS. 3a to 3 e;

FIGS. 5a to 5c are front views of initial steps of a build sequence of atubular component using the system according to another preferredembodiment of the present invention, with an alternate distribution ofremovable inserts with respect to the molds, with no insert on a firstmold, one (1) insert on a second mold and two (2) inserts on a thirdmold;

FIGS. 6a to 6c are perspective views of the build sequence of thetubular component shown in FIGS. 5a to 5 c;

FIG. 7 is a perspective view of a build sequence of a fuselage componentusing the system according to another preferred embodiment of thepresent invention and showing installation of composite layup materialsby personnel; and

FIGS. 8a and 8b are schematic views of a stepped-lap joint interface anda scarf-joint interface respectively

PREFERRED EMBODIMENTS OF THE PRESENT INVENTION

An object of the present invention is to manufacture a compositematerial assembly, such as, but not limited to, a tubular profilestructure from two or more longitudinal section components. The wholeassembly can be cured in one step in order to form a one-piece tubularstructure, such as, for example, a fuselage. Hence, the components thatwill constitute the whole assembly are joined before curing occurs andthen the whole assembly is cured through co-curing of these components,producing an end product without any overly apparent seams.

Referring to FIGS. a to 1 c, according to a first preferred embodimentof the present invention, a system 10 for fabricating a compositematerial assembly is disclosed. The system 10 includes a first mold 12for receiving a first module 13 made of composite material. The firstmold 12 has a first composite material laminate support structure 14having first and second opposite edges 16, 18. The first mold 12 alsohas a first attachment interface 20 for attachment of the first mold 12to an adjacent mold 22. The system 10 also comprises a second mold 22for receiving a second module 23 made of composite material. The secondmold 22 includes a second composite material laminate support structure24 having first and second opposite edges 26, 28. The second mold 22also has a second attachment interface 30 for attachment of the secondmold 22 to the first mold 12.

The system 10 further comprises a removable insert 32 extending beyondthe second edge 28 of the second mold 22. The insert 32 is shaped suchthat it would contact the first mold 12 if the first and second molds12, 22 were attached together and would prevent attachment therebetweenif the insert 32 was present.

The first module 13 comprises a first laminate 34 covering the firstlaminate support structure 14. The second module 23 comprises a secondlaminate 36 covering the second laminate support structure 24 andextending over the removable insert 32. As better shown in thetransition between FIG. 1a and FIG. 1b , the removable insert 32 isremoved from the second mold 22 prior to assembly of the first mold 12to the second mold 22. As better shown in FIG. 1b , a section 38 of thesecond laminate 36 extending over the removable insert 32 overlaps overthe first laminate 34 after closing and assembly of the first mold 12onto the second mold 22. At the initial closing of the molds 12, 22, thelaminates 34, 36 are not cured or solidified, allowing the requiredtackiness, softness and flexibility to ensure proper intermesh and layupat the interface.

Preferably, the first laminate 34 has a first interface profile 40shaped to fit into a complementary second interface profile 42 of thesection of the second laminate 36 extending over the removable insert32. The second laminate 36 can therefore overlap over the first laminate34, and form a joint without an overly apparent seam, between the firstand second modules 13, 23. Preferably, the interface profiles 40, 42 arechosen to form a stepped-lap joint. During initial placement of thelaminates 34, 36 on the molds 12, 22, the removable insert 32 providesthe extension surface that is required to form the stepped-lap joint.

Preferably, when the composite material assembly is a tubular component,the final assembly results from two or more joints. Given that thechosen type of joint for this application must minimize any overthickness in order to obtain a uniform structure thickness along theperimeter of circumference of the assembly, it is preferable to use atype of joint that requires superimposing two half-elements, preferablythrough a stepped-lap interface as mentioned above. In other embodimentsof the present invention for fuselage applications, the stepped-lapinterface, as shown in FIG. 8A, at the sectional interface betweenmodules could be changed for a scarf-type lamination, as shown in FIG.8B without affecting the required constant thickness of the fuselagestructure.

As shown in FIG. 1b , manufacturing of a joint in accordance with thepresent invention requires that one section 38 of the laminate 36overhangs temporarily and is therefore not supported beyond the edge 38of the mold 32. This overhanging configuration is required for theperiod of time between removal of the insert 32 and closing of the molds12, 22 for forming the assembly.

In order to allow closing of the molds 12, 22, each of the half-elementsof the complementary interface profiles 40, 42 to be stacked must avoidcontact with each other during the closing movement of the molds 12, 22,as there can be a risk of localized pre-adherence, before the twohalf-elements are positioned correctly. Any incorrect positioning of thetwo sides of the interface for the laminate could result in theformation of air pockets and result in an abnormal discontinuity in thestructural laminate in the joint assembly zone.

In order to avoid this possibility of pre-adherence between the twohalf-elements of the joint prior to the final closed position of themolds 12, 22, the removable insert 32 preferably has a geometrical formshaped to position the overhanging section 38 of the laminate 36, withthe interface profile 42, above its corresponding interface profile 40on the other mold 12 without incurring any contact or pre-adherence,after the insert 32 is removed.

Preferably, the surface of the insert 32 on which the overhangingsection 38 of the laminated interface is resting has an angular positionof at least 10° and preferably between 10° and 15° with respect to atangential direction of the second laminate 36 of the second mold 22, atthe edge 28 of the second mold 22 where the removable insert 32 ispositioned, towards an inner side of the second mold 22.

For fuselage applications, the required laminate construction for thefuselage can be a monolithic configuration, as shown in FIGS. 1a to 1c ,or sandwich/core structure, as shown in FIGS. 2a to 2c , or acombination of the two. Preferably, in the case of a fuselagesandwich/core structure, as shown in FIGS. 2a to 2c , the initiallaminated assembly produced with the removable insert 32 and the molds12, 22 is the same but this laminated assembly may now be designated asan “outer skin” 74 a. As shown in FIG. 2b , the “outer skin” 74 areceives a sandwich honeycomb core 72 followed by an “inner skin” 74 bwhich could be of different construction. The inner skin laminate 74 bis terminated also at its longitudinal edges 78, 79 by a stepped-lapgeometry being in full contact with the sandwich core 72 surface.Preferably, the laminates are made of “out of autoclave” carbon-epoxypre-preg and the sandwich/core structure comprises a Nomex™ honeycombcore, however, other materials may be used.

Preferably, the removable insert 32 is a structural element. However,the removable insert may be an inflatable structure, or any otherretractable molding structure known to a person skilled in the art.

Preferably, the attachment interfaces 20, 30 are hinge-type interfaces.However, other types of attachment interfaces may be used. Moreover, theattachment interfaces 20, 30 may comprise a cam assembly in order toprovide a sufficient amount of clearance for the overhanging section 38of the laminate 36 to avoid inadvertent contact and pre-adherence withthe other side of the interface.

Preferably, the system 10 further comprises a flexible elastomeric sealat a joint interface between the first and second molds 12, 22. Theflexible elastomeric seal provides vacuum integrity of the mold assemblyneeded for the curing procedure.

Preferably, a release agent is applied to the first and second molds 12,22 prior to laying down of the first and second modules 13, 23 thereon.The release agent is preferably one of three types: (i) liquid or paste,(ii) in the form of a plastic film and (iii) of a permanent type such asa Teflon™ coating and one skilled in the art can select the appropriateone for its particular need. Additionally, other types of release agentsmay be considered, The release agent is applied in each mold to allowremolding of other modules after a curing step.

Preferably, the first and second molds 12, 22 are portions of acylindrical structure. The system can therefore be used to form a curvedassembly as shown in FIGS. 3a to 3e and FIGS. 4a to 4e . Preferably, themolds have a geometric shape adapted to form a tubular-profiledstructure and comprises at least two 180° sections or preferably three120 sections.

Referring to FIG. 3a , when the molds are made of three sections, acentral mold 12 rests on the ground with the two other molds 22, 52placed adjacently.

Preferably, when the assembly molds 12, 22, 52 comprise three sectionsto form a cylindrical structure, the removable inserts can be positionedin different manners. In a preferred embodiment of the presentinvention, one insert is associated with each mold 12, 22, 52, as shownin FIGS. 3a to 3e , since such a configuration allows for themanufacture of three (3) identical molds/inserts. However, otherconfigurations can be considered. For example, no insert can beassociated with the first mold 12, one insert can be associated with thesecond mold 22, and two inserts can be associated with the third mold52, as shown in FIGS. 5a to 5 c.

The closing sequence of the different molds 12, 22, 52 is not influencedby the positioning and distribution of the inserts among the differentmolds because a clearance zone has been designed into the shape of themolds in order to position, within this clearance zone, the overhangingsection 38 of the laminate 36 to avoid contact between the two sides ofthe interface of the assembled laminate interface during closing of themolds.

Preferably, for assembly of cylindrical fuselage components, among otherapplications, three molds 12, 22, 52 are provided. As better shown inFIGS. 5a to 5c , the system 10 for fabricating a composite materialassembly comprises a third mold 52 for receiving a third module made ofcomposite material. The third mold 52 includes a third compositematerial laminate support structure 54 having first and second oppositeedges 56, 58. The third mold also has a pair of opposite third andfourth attachment interfaces 60, 62 for attachment of the third mold 52to the first and second molds 12, 22. The third mold 52 also has secondand third removable inserts 64, 65 extending beyond the first and secondedges 56, 58 of the third composite material laminate support structure54. The first mold 12 comprises a fifth attachment interface 66 forattachment of the first mold 12 to the third mold 52. The second mold 22comprises a sixth attachment interface 68 for attachment of the secondmold 22 to the third mold 52. A third laminate 70 covers the third layupstructure 54 and extends over the second and third removable inserts 64,65. The first, second and third removable inserts 32, 64, 65 are removedfrom the second and third molds 22, 52 prior to assembly of the first,second and third molds 12, 22, 52. A section 71 of the third laminateextending over the second removable insert 64 overlaps over the firstlaminate 34 after closing and assembly of the third mold 52 onto thefirst mold 12. Another section 73 of the third laminate extending overthe third removable insert 65 overlaps over the second laminate 36 afterclosing and assembly of the third mold 52 onto the second mold 22. Asmentioned above, the distribution of the inserts 32, 64, 65 among thedifferent molds as shown in FIGS. 5a to 5e may vary for a selectedassembly closure sequence and correspond, for example, to thedistribution of inserts 32, 64, 65 shown in FIGS. 3a to 3e . In FIGS. 5ato 5e , the distribution of the inserts is such that first mold 12resting on the ground has no inserts and sections 71, 38 overlap overthe first laminate which may be a more practical sequence of assembly ofthe laminates in certain assembly configurations.

According to the present invention, there is also provided a method forfabricating a composite material assembly comprising the steps of:

-   -   a) providing an assembly system 10, as shown in FIGS. 1a to 1c        comprising:        -   a first mold 12 for receiving a first module 13 made of            composite material, the first mold 12 comprising:            -   a first composite material laminate support structure 14                having first and second opposite edges 16, 18; and            -   a first attachment interface 20 for attachment of the                first mold 12 to an adjacent mold 22; and        -   a second mold 22 for receiving a second module 23 made of            composite material, the second mold 22 comprising:            -   a second composite material laminate support structure                24 having first and second opposite edges 26, 28;            -   a second attachment interface 30 for attachment of the                second mold 22 to the first mold 12; and            -   at least one removable insert 32 extending beyond the                edge 28;    -   b) laying down the first module 13 on the first mold 12, the        first module 13 comprising a first laminate 34 covering the        first laminate support structure 14;    -   c) laying down the second module 23 on the second mold 22, the        second module 23 comprising a second laminate 36 covering the        second laminate support structure 24 and extending over the        removable insert 32;    -   d) removing the removable insert 32 from the second mold 22;    -   e) assembling the first mold 12 with the second mold 22 while        overlapping a section 38 of the second laminate 36 extending        over the removable insert 32 over the first laminate 34.

Preferably, the method further comprises the step of f) curing theassembled first and second modules 13, 23 in an oven. When the methodaccording to the present invention is used to manufacture a fuselageassembly, considering the fact that the entire composite structure ofthe fuselage has been realized in a complete uncured state and that thecomposite structure is fully assembled in a tubular profile, the entirefuselage assembly inside the closed mold has to be solidified by puttingit under vacuum and heat inside a curing oven. Under only one “heat andpressure cycle” the pre-preg laminate and adhesive will cure andsolidify to generate a one-piece tubular section of fuselage without anoverly apparent seam. It is however understood by one skilled in the artthat any appropriate curing process is possible pursuant to theinvention.

Preferably, the one-piece section of fuselage produced using the systemor method may integrate or comprise floor attachment members, a cockpitwindshield, cabin windows and passenger door surrounding structures. Allof these features may be all cured in one step only. The system andmethod according to the present invention can be used for any portion ofa flying vehicle which possesses a tubular profile with a need to beco-cured for reducing any overly apparent seam, such as any cabin of anaircraft.

Referring to FIG. 7, the system and method according to the presentinvention can be used for manufacturing of one-piece fuselage sectionsand facilitate the layup of composite pre-prep material on the molds 12,22, 52 in an almost horizontal position, thus reducing the countereffect of gravity when compared to a tubular or cylindrical molds.

Although preferred embodiments of the present invention have beendescribed in detail herein and illustrated in the accompanying drawings,it is to be understood that the invention is not limited to theseprecise embodiments and that various changes and modifications may beeffected therein without departing from the scope or spirit of thepresent invention.

1.-11. (canceled)
 12. A method for fabricating a composite materialassembly comprising the steps of: a) providing an assembly systemcomprising: a first mold for receiving a first module made of compositematerial, said first mold comprising: a first composite materiallaminate support structure having first and second opposite edges; and afirst attachment interface for attachment of the first mold to anadjacent mold; and a second mold for receiving a second module made ofcomposite material, said second mold comprising: a second compositematerial laminate support structure having first and second oppositeedges; a second attachment interface for attachment of the second moldto the first mold; and at least one removable insert extending beyond atleast one of said first and second edges of the second mold; b) layingdown the first module on the first mold, the first module comprising afirst laminate covering the first laminate support structure; c) layingdown the second module on the second mold, the second module comprisinga second laminate covering the second laminate support structure andextending over the at least one removable insert; d) removing the atleast one removable insert from the second mold; e) assembling the firstmold with the second mold while overlapping a section of the secondlaminate extending over the at least one removable insert over the firstlaminate.
 13. The method according to claim 12, wherein the first andsecond molds are portions of a cylindrical structure.
 14. The methodaccording to claim 12, wherein the first laminate has a first interfaceprofile shaped to fit into a complementary second interface profile ofthe section of the second laminate extending over the at least oneremovable insert and overlapping over the first laminate, for forming ajoint between the first and second modules.
 15. The method according toclaim 12, wherein the at least one removable insert comprises a laminateoverhang support surface, said laminate overhang support surface beingoriented at an offset angle of at least 10° with respect to a tangentialdirection of the second laminate of the second mold, at the at least oneof said first and second edges of the second mold where the at least oneremovable insert is positioned, towards an inner side of the secondmold.
 16. The method according to claim 15, wherein the offset angle isbetween 10° and 15°.
 17. The method according to claim 12, wherein theassembly system further comprises a flexible elastomeric seal at a jointinterface between the first and second molds.
 18. The method accordingto claim 12, further comprising the step of, prior to step b), applyinga release agent to the first and second molds prior to layup of thefirst and second modules thereon.
 19. The method according to claim 12,further comprising the step of f) curing the assembled first and secondmodules.
 20. An aircraft fuselage comprising a composite materialassembly fabricated according to claim
 12. 21. The aircraft fuselageaccording to claim 20, wherein the fuselage is a solid laminate.
 22. Theaircraft fuselage according to claim 20, wherein the fuselage is asandwich structure.
 23. The aircraft fuselage according to claim 20,wherein the fuselage is a combination of a solid laminate in somelocations and a sandwich structure in other locations.
 24. The aircraftfuselage according to claim 20, comprising at least one componentselected from the group comprising floor attachments, cockpitwindshields, cabin windows and passenger door surrounding structures.